Jet propulsive devices



Nov. 22, 1955 T. D. GREGG JET PROPULSIVEDEVICES 4 Sheets-Sheet 1 FiledOct. 25, 19.50

Nov. 22, 1955 T. D. GREGG JET PROPULSIVE DEVICES 4 Sheets-Sheet 2 FiledOct. 25, 1950 INVE TOR..

Nov. 22, 1955 Filed Oct. 25, 1950 ENTHALP: 5.711 PER 1.5

T. D. GREGG JET PROPULSIVE DEVICES 4 Sheets-Sheet 3 6 ENT'ROPY FIG. 8

IN NT R.

1955 T. D. GREGG 2,724,31

JET PROPULSIVE DEVICES Filed Oct. 25, 1950 4 Sheets-Sheet 4 2,724,238 vJET PROPULSIVE DEVICES Tresham D. Gregg, New York, N. Y.

Application October 25, 1950, Serial No. 191,979

6 Claims. 01. 60--35.6)

Thisinvention relates to jet propulsive devices and particularly tomeans of producing such propulsive reaction by a jet from a gas turbineengine having a nozzle whose area of cross-section may be varied at willin conjunction with a segmentally-adjustable aerofoil ring thrustaugmentor.

These and other objects of thisinvention are now summarized as follows:3

1. In typical jet propulsion engines such as the Goblin II describedinfthe issue of February 24, 1946 of the magazine Flight, the tail pipeor propulsion nozzle is of fixed area. This fact limits the range ofthrust economically obtainable from the engine and also limits itspropulsive efficiency at speeds less than400 miles per hour. Thisinvention increases such range of thrust and also increases thepropulsive efficiency of the engine at moderate speeds by providing theengine with a segmentally-adjustable nozzle or tailpipe having a widerange of area variation. This is a tube consisting of a plurality ofcyliudricallyor comically-curved plates over-lappingly assembled andmovable radially with respect to the axis of the nozzle, beingmaintained in close contact at the overlaps by the pressure of the fluidwithin the tube.

2. The principal loss of energy and thrust in all fluid reactionpropulsive devices now in use is theso-called leaving loss, that is, thekinetic energyof the propulsive jet as it leaves theprojecting nozzle.One of the objects of this invention, therefore, is to reduce to aminimum the leaving loss of the propulsive jet from a gas turbine jetengine. This object is bestaecomplished by adding to the propulsionsystem an aerofoil ring, preferably one that is segmentally-adjustable.Such a device utilizes the otherwise wasted kinetic energy of thedriving jet as it leaves the nozzle by entraining and acceleratingthrough the ring a large volume of outside airwhich, as it passesthrough the ring, mixes with the hot gas of the jet, cools it, reducesits speed, and at the sametime increases the flow of momentum of thesystem and henceits thrustybut without at the same time increasing thefuel consumption of the engine. a

3. In order to make effective the addition of such an aerofoil ring tothe system, the stream of outside air thus drawn into the aerofoil ringduct should be guided smoothly into contact with the driving jet at thecritical point in the ring channel for maximum thrust augmentation,through an annular duct having no sudden variations of area andadaptable for all possible nozzle diameters. The invention accomplishesthis by attaching to that portion of the nozzle which projects beyondthe nacelle or engine fuselage a segmentally adjustable conoidal sur-2,724,238 Patented Nov, 22 1955 the inducted air must pass through anair gap between the fuselage and the leading edge of the aerofoil ringin a direction normal to the ring axis. Such an arrangement makes theratio of the upstream reaction or thrust boost of the ring to the thrustof the jet substantially independent of the aeroplane speed. i

5. Since the aerofoil ring when in efiicient operation can deliver anindependent thrust of 40% or more of the normal jet thrust at allspeeds, it must be supported in the correct position with respect to theaeroplane under all conditions of speed and direction. This isaccomplished by connecting the leading edge of the aerofoil ring to thenacelle or fuselage by means of a plurality of streamlined struts orlinks passing through close-fitting slots in the nacelle or fuselage andstrong enough not only to transmit the ring reaction to the aeroplanebut also to resist all stresses due to the weight of the augmentor andto the bending moments and shears and tortional forces due to suddenchanges in direction of the aeroplane in flight. These struts or linksmust also supply means of controlling the width of the air gap and thussupply the means of making quick changes in thrust such as arefrequently needed during landing and take-off, without requiringcorresponding changes in the engine throttle with the consequent dangerof killing the engine at a critical time. This device also serves theimportant function of giving the pilot a continuous measure of the netaerofoil ring thrust. The method of support also permits the necessaryexpansion and contraction of the ring.

6. All bodies moving through a fluid, whether streamlined or not, aresubject to the drag of skin friction. Skin friction is proportional inintensity to the thickness of the boundary layer, a thin layer of airwhich virtually attaches itself to the surface of the moving body. Theboundary layer begins to form at the forward end of the body andgradually increases in thickness toward the rear. Friction drag can bematerially reduced by causing the boundary layer to be sucked into theinterior of the body at intervals over its surface. The air gap betweenthe nacelle and the aerofoil ring permits this sucking in of theboundary layer that has formed upstream so that the skin friction dragintensity on the outside surface of the ring is thereby materiallyreduced.

7. It is an important object of one form of this'invention to makepossible a tail pipe or nozzle of variable discharge area having a shapeand length suitable for after burning either with or without theaddition to the system of an aerofoil ring thrust augmentor. Afterburning, as the name implies, is the mixing of new fuel with the turbineexhaust gas in the tail pipe to be there ignited and burnt at a veryhigh temperature, the greatly expanded gas being then discharged throughan enlarged nozzle at correspondingly increased velocity and augmentedthrust. Before igniting this gas mixture two things are essential,namely, the tail pipe channel must be so shaped as to diffuse orrecompress the gas after it has passed through the turbine and, second,a device or devices must be inserted into the channel after the fuel hasbeen introduced to mix it thoroughly with the exhaust gas before it isignited. Since such devices materially reduce the: thrust, especiallyfor normal operation, they should be automatically removable afterperforming their function. These objects are accomplished in thisinvention first by forming the whole tail pipe from a plurality ofoverlapping segments movable radially in the same manner as theadjustable nozzle above described, second, by giving the pipe assemblythe longitudinal variations in diameter required for diffusing the gas,and finally by utilizing the radial movement of the component segmentsby a simple mechanism to raise and lower a plurality of hinged platescapable of stirring up the passing gas during after burn} ing and .of.being folded .out of .the .channel .for normal operation.

These and other objects of'this invention will be understood when thefollowing description is read in connection with :the attached drawings.

Fig. :1 .showsaside elevationof the engine nacelle and the attachedaerofoil ring thrust augmentor.

Fig. 12 shows the venginewwith its nacelle partly broken away, theadjustable nozzle in a contracted position in axial section togetherwith a portion of the forward end :of the aerofoil ring, also in acontracted position.

Fig. 2a shows a part of the rear end of the engine nacelle with aportion of the nozzle and of the forward part of the :aerofoil ring,both in an expanded position.

LF-ig. 2b shows in enlarged section a portion of the gas turbine withparts of the inner cone or bullet and its supporting struts togetherwith the conical part, in seciion, of the segmentally adjustable tailpipe.

Figs. 3 and 3a .show partial sections C--C and BB of the segmentallyadjustable fairing cone F and the after-face of the nacelle.

Fig. 4 shows a portion of section A--A of the tailpipe with itsadjusting engines in the contracted position.

Fig. 5 shows a cross-section of the cylinder of one of the engines whichsupports the aerofoil ring and controls the-air gap.

Fig. 6 shows an .enlarged partial elevation of the conoidal initialportion of the adjustable nozzle with its hinged component members andone dilating or adjusting engine.

Fig. 7 shows a partial cross section and upstream axial view D--D of theinitial portion of the tail pipe in both contracted and dilatedpositions, shown in elevation in Fig. 6.

Fig. 8 shows a Mollier or enthalpy-entropy diagram defining graphicallythe characteristic changes in enthalpy, pressure and entropy which takeplace in a typical turbo jet engine alone and when combined with anaerofoil ring thrust augmentor both in the normal and 'afteriburning" orreheat cycles.

Fig. 9 shows the combination of a tail pipe arranged and. proportionedfor after burning in conjunction with an aerofoil ring thrust augmentoras in the second embodiment of this invention.

Fig. 10 shows in enlarged detail a partial transverse section of thetail pipe with the stabilizers and their operat-ing' mechanism both inthe after burning and normal position.

Figs. 11, 12 and 13 are enlarged details of tangentially operatingpistonengines for use if space is too small for the radial engines, shown inpartial cross-section, both in open and closed position, and inelevation, of a segmentally-adjustable structure composed of overlappingplates.

Fig. 2 shows in sectional form a normal gas turbine engine having anadjustable tail pipe or nozzle in conjunction with an aerofoil ringthrust augmentor partially shown, also in section, which forms the firstembodiment of the present invention.

In Fig. 2, 6 represents one of the air intake orifices which areconnected by expanding ducts to a turbocompressor, as 7. The compressoris connected to sixteen burners, two of which, designated 8, are shown.These are joined at their downstream ends to a nozzle ring casting 24shown in enlarged section in Fig. 2b. This casting is supportedpartially by struts 26 and partially by the cone 18. This cone alsosupports thedownstream bearing'of the hollowshaft 12 which carries thegas turbine rotor 11. A plurality of stream-lined struts 25, integralwith the nozzle ring casting 24, support the cone or bullet 21 whichforms the inner wall of the annular channel 20. The outer wall 22 ofthis channel, formed of a plurality of overlapping conoidally curvedplates is connected by hinges 23, Fig. 2b, to the nozzle ring casting24. The downstream ends of plates 22 are slidingly connected by means ofthe slotted double hinges .30., Figs. .2 and 6, to correspondingsegmental plates forming the discharge nozzle 27. Each of these plateshas a longitudinal stiffener 31 to which are attached the piston rods oftwo independently controlled sets of radially operating double-actingpiston engines 32 and 32a, shown also in Figs. 4 and 6 with two fluidsupply manifolds 49 and manifold feed pipe 50, supported by acylindrical plate or drum 33 which bears against a plurality of blocks17 fastened to the engine nacelle 19. This nacelle is squared with the.engineaxis at its downstream end and has a rounded corner, forming awindshield 1%, Figs.

2, 2a and 3a for the aerofoil ring thrust augmentor 2,

Figs. 1, 2 and 2a, which is held in position by a plurality of struts orlinks 3 attached by pins 16, Figs. 2 and 2a, to the leading edge of thering and by pins 15, Figs. 2 and 2a, to the piston rods 13 of acorresponding number of double-acting hydraulic engines 14, Figs. 2, 2a,4 and 6, attached to the nacelle 19. Slots 1%, Figs. 3 and 3a, areprovided in the corner of the nacelle and in the blocks supporting thedrum '33 to permit the struts 3 to move radially as well aslongitudinally and fitting the struts or links 3 closely so as toprevent any twisting motion of the aerofoil ring which they support. Twofluid supply pipes .51 provide the driving fluid for the said engines14.

The outer or downstream portion of the nozzle extends beyond the squaredend of the nacelle to the critical section of the aerofoil ring duct,called the aerodynamic center of the" ring, the position of this centerbeing fully described in my Patent No. 2,475,022, column 3, lines 44 to51, and column 7, lines 7 to 30. To each of the segments of the nozzle27, Figs. 2, 2n, 3 and 3a, is fastened a radial diaphragm 36 which has acurved outer edge to which is fastened a conoidally curved fairing plate35 extending from the, nacelle to the nozzle discharge orifice. Theseplates when overlappingly assembled form a smooth fairing cone F, asshown enlarged in Figs. 3 and 3a, integral with the diaphragms 36 andthe nozzle plates 27, and bear slid-ingl'y, through an anti-frictiondevice if required, upon the squared rear end of the nacelle 19b, and 1Fig. 2a shows the nozzle fairing cone assembly F in its dilated orexpanded position in relation to the forward end of the aerofoil ringcorrespondingly expanded as for maximum thrust, which forms, with theinner surface of the ring, the annular channel 28. In each of Figs. 1, 2and 2a the variable air gap 4 between the engine nacelle at and theleading edge of the aerofoil ring is shown As above indicated, a secondembodiment of this invention consists of a gas turbine engine with asegmentally adjustable tailpipe suitably shaped to serve in expandedposition as an after burner in conjunction with an aerofoil ring thrustaugmentor; and also in contracted position as a normaljet projector.Fig. 9 shows such an arrangement. The parts 21', 22, 23, .24, 25,26- and30 are also shown in-Figsy2, 2b and 6. 20a is' anannular channel such as20 of Figs. 2 and 2b but greatly expanded. In Fig. 9, channel 37 isformed of a plurality of overlapping conically curvedsegments connectedto the corresponding. segments of channel20a by the slotted doublehinges 30, Figs. 6 and 7-. The segments 37 are joined integrally as bywelding or other means atthe point of maximum expansion withcorresponding convergent-divergent segments of channel 370, thedivergent-portion of which extends with its attached fairing coneFbeyond the nacelle into the duct of the aerofoil ring 2, hereshownexpanded for maximum thrust augmentation. At or near the junction of- 37and 37a are located a plurality of-hinged plates or stabilizers 38 witharms 39;projecting outsidethe pipe. These plates which project into thepipe channel during after burning as shown in Fig. 9 and in enlargedpartial cross-section; in Fig. 10 stir up or mix the'passing gas aspreviously mentioned. When the tail-pipecontracts for normal operationas shown by the broken lines in-Figs. 9 and 10, these stabilizers arefolded back to lie fiat against the Wall of the pipe by means oftheoutside projecting arms 39"wliiciieon'nectb'y means of universal ofblocks 43 attached to the nacelle.

. the gas.

joints 40 to other; arms 39!: which are held bysimilar joints 40 tosupport castings 41 resting uponacircular plate or drum 42. This drum is,carried on a-plurality The tail pipe 37.--37a issupported and movedradially by piston engines arranged in two independently operating sets32 and 32a carried on drums, as 33, which arein turn supported on blocks44 and 45 the necessary driving fluid being conveyed to said engines bymanifolds49 and manifold supply pipes 50. The outboard section .of thenozzle carries the fairing cone F as. previously described bearingslidingly on the after-face of the nacelle at 19b, Figs. 2a, 3 and 3b. eI

.An alternative device for effecting the circumferential expansion andcontraction of the nozzle ducts of this .invention, or any structurecomposed of overlapping plates, where space is too small for radialengines, is shown contracted in Fig. 11 and expanded in Fig. 12, thetangentially-operating. piston engines 46 being connected to theadjacent overlapping plates by means of pins 47 and blocks 48. s Theblocks 48 serve as stops to limit contraction and the piston cylindersserve as stops to. limit expansion. Fig. 13 is an outside elevation ofthree of these engines and attached plates. h

, In the operation of this invention in conjunction with a turbo-jetengine such as the Goblin II. air normally at high speed relative to theplane enters two intake orifices, one of which is shown in Figs. land 2M6, at atmospheric pressure and heat content Po and Ho respec tively onthe diagram Fig. 8. It passes through ducts to a compressor as 7, Fig.2, which it enters at a pressure P1 called the ram pressure" with a heatcontent or enthalpy of H1 in B. t. u. per lb. of gas as noted on thediagram, Fig. 8. In the compressor it is raised in pressure and enthalpyto P2 lb. per square inch and H2 units of enthalpy and passes thenceinto 16 combustion chambers of which 2 are shown as at 8, Fig. 2, intowhich liquid fuel is fed from a manifold 9 which mixes with a thecompressed air and is ignited, burning at a high temperature, its totalheat represented by H3 and at a pressure P3. The gas thus producedexpands continuously through turbine nozzles formed of aerofoil shapedvanes or blades, Fig. 2b at 10, in the nozzle casting 24, against themoving blades of the turbine rotor Fig. 211,11. The hot gas leaves theturbine at pressure P4 and enthalpy H4, Fig. 8, whence it continues toexpand through the annular duct 20, Figs. 2 and 2b, formed by the,bullet 21 and the converging outer cone 22, and thence through the duct27 to its exit in the central duct of the aerofoil ring where it has thepressure P6 and heat content He.

In the case of a normal jet engine, such as the Goblin II without theaerofoil ring, the discharge pressure would be asmospheric, P0 and theenthalpy H5, Fig. 8, the heat H5-Ho B. t. 11. per lb. of gas beingrejected in this cycle of operation. This rejected heat is proportionalto the leaving loss or kinetic energy of the discharged gas. In thefirst embodiment of the present invention, however, the dischargepressure P6 and enthalpy H0 at the nozzle exit do not represent thefinal state of After leaving the nozzle the gas continues to do usefulwork by drawing into the aerofoil ring through the air gap 4 and theannular duct 28, Figs. 1 and 2, a

much larger weight of outside air by reason of the pressure drop PO-PG;and accelerating it and compressing it to atmospheric pressure at thetrailing edge orifice 29. When this air passes the exit orifice of thenozzle 27 it has the pressure P6 and enthalpy H7, Fig. 8. Between thispoint and the trailing edge orifice of the aerofoil ring 29, Fig. 1, itis mixed with the hot gas from the driving nozzle 27 and the mixture isdischarged through the trailing edge orifice 29 at atmospheric pressurePo and enthalpy Ha as shown in Fig. 8. The lost heat Ha-Ho per lb. ofdriving gas is seen to be less than in the case of the normal jet engineas above described, and a consequent reduction .in the leaving loss witha correspondingincrease in thrust is thus achieved by the firstembodiment of this invention.

In the operation of the after burner which constitutes a secondembodiment of the invention, the gas leaving the turbine passes into agradually enlarging annular duct at 20a, Fig. 9, formed by the bullet 21and the same segmentally hinged conoidal surface as 22, Figs. 2 and 6,in the first embodiment above described, each of the component segmentsof which is also attached by means of slotted double hinges as 30, Figs.2, 6 and 7, to corresponding segments of the after burner tail pipe, thefirst section of which, 37, continues to increase in diameter.

Between the hinged joint and the point of its maximum diameter liquidfuel is sprayed into duct 37, by means not shown. At or near the pointof maximum diameter where 37 joins the convergent-divergent section 37a,Fig. 9, the, stabilizer plates 38, Fig. 9, stir up the passing gas tocause the fuel to mix thoroughly with it. The mixture is ignited andburns at a temperature generally very much higher than thatof the gaswhich entered the turbine, with maximum enthalpy being H'4 and pressureP4. and expands finally to atmospheric pressure P0 with a heat contentH's for the jet engine without the aerofoil ring, rejecting the heatH's-Ho, and to P's and H's at the nozzle exit when such ring is present,whence it mixes with the outside air drawn into the ring as beforedescribed, the whole being discharged from the ring orifice 29, Fig. 1,at atmospheric pressure P0 and heat content H's. Again it is apparentthat the heat given up to the atmosphere, H's-H0, i. =e., the leavingloss, is less than that of the after burning jet=engine Without theaerofoil ring, i. e., H's-Ho. Hence the statements in the secondparagraph of this specification are justified for both of theembodiments of the present invention.

It is obvious that the thrust of a nozzle propelled jet is impressedupon the nozzle walls by the jet, and by the nozzle upon the engine andits nacelle, either directly or indirectly. It is understood, therefore,that the method and means of transfer of jet thrust to the enginenacelle and aeroplane must be adequate and not necessarily those shownand described in the specification.

The speed of the. driving jet is normally equal to or greater than thatof sound. The jet therefore has great lateral stiffness due to itsmomentum and, together with the annular stream of inducted airsurrounding it which also has great momentum, forms an elastic supportto the aerofoil ring against sudden lateral movement and shocksoccurring during flight.

While the specific means for applying pressure to the various pistonengines described herein have been shown on the drawings, it is to beunderstood that any suitable means, either manually or automaticallycontrollable, may be employed, such as those shown and described in myPatent No. 2,475,022 for tangentially controlling a segmentallyadjustable aerofoil ring.

As at present constructed, tail pipes for after burning" andparticularly the means now in use for varying the area of discharge aresubject to severe temperature stresses due to the great variation intemperature that occurs between the entrance to and the exit from thetail pipe; and also between normal and after burning operations.Moreover, the devices now in use for varying the area of discharge aresubject to the danger of over confining explosive gas mixtures. Thepresent invention, because of the segmental character of the tail pipe,readily adjusts itself to unequal expansion and contraction due to largetemperature differences, and the discharge orifice gives free egress toexploding gas.

As indicated in paragraph numbered 1 in column 1 it is assumed that thesurfaces at the points of contact between the overlapping componentsegments of the adjustable nozzle or tail pipe will be smooth andaccurately manufactured so that the gas pressure will make the juncturesgas-tight.

Pipes for conveying cooling air from the compressor to ,7 the hollowturbine shaft may be passed through the struts 25 which support theinner cone as in the Goblin 'II engine.

.I claim:

1. A gas turbine jet aeroplane engine in combination withan aero'foilring thrust augmentor, said engine having a nozzle and also havingastream lined body immediately in front of the aerofoil ring said bodybeing substantially larger in diameter thanjthe outer diameter of the'aerofoil ri'ng, said body having a plane surface presented toward theleading'edge of the aerofoil ring and perpendicular to the axis of thesaid aerofoil ring, said nozzle being substantially smaller than theminimum diameter of the ring duct andof such length that its downstreamend is positioned "at the aerodynamic center of the ring.

2. An airplane power plant including, in combination, a gas turbine ,jetengine, a nacelle for said engine having a squared after face, a jetpropelling nozzle for said engine having a conoidal fairing extendingfrom its orifice to the square after face of the said nacelle, anaerofoil ring thrust augmentor of segmentally variable diameter coaxialwith said nozzle and means to fasten said augmentor to said engine, saidfastening means being arranged to vary the width of the air gap betweenthe leading edge of said augmentor and the after face of said nacelle tohold the variable aerodynamic center of said augmentor at the plane ofthe nozzle discharge orifice and to permit varying the diameter andprofile of said augmentor.

3. In combination, a gas turbine jet aeroplane engine positioned in 'ahollow, stream-lined nacelle having a nozzle projecting axiallydownstream into. the duct of an aerofoil ring thrust augmentor, the reardiameter of said nacelle being substantially larger than the -maximumdiameter of the said augmentor, the outer portion of the rear face ofsaid nacelle being square with its axis, said nozzle forming a largeannular duct with the inner surface of the augmentor, the entrance tosaid annular duct being defined by the squared rear face of the saidnacelle and the leading edge of the aerofoil ring thrust augmentoradapted to be contracted and expanded; a plurality of, struts adjustablyconnecting said augmentor to the said nacelle and each arranged to varyits angular position toward the axis of the said augmentor at itsforward end during contraction and expansion of said augmentor.

4. The combination with a gas turbine jet airplane engine having anozzle the diameter of which may be varied of a fairing cone surroundingsaid nozzle and expandible and contractible therewith an aerofoil ringthrust augmentor into which the nozzle extends, the downstream end ofsaid nozzle being substantially smaller than the minimum diameter of theaugmentor duct and located substantially at the aerodynamic center ofthe said ring section, a nacelle for said engine the squared after faceof 8 which ;is in sliding contact with the fairing cone of the noz: 21cand is substantially larger in diameter than the leading edge of theaerofoil ring, means to attach saidaugmentor to said engine comprising aplurality of struts, pivotally fastened at one end to the leading edgeof said augmentor and passing through close fitting slots in the enginenacelle to pivots in the ends of the piston rods of a correspondingnumber 'of double acting hydraulic piston engines within the nacelle,the direction of motion of said piston being parallel to the jet'engine'axis to control the width of the air gap between thelea-ding'edge of the aerofo il ring and the engine nacelle.

5. A gas turbine jet engine including, in combination, a nozzle ofvariable cross-section composed of a plura'lit'y of overlapping segmentsprojecting beyond the engine nacelle, the projecting portion of thecomponent overlapping segments being connected by radial diaphragmsto acorresponding plurality of conoidally curved overlapping segmentsforming an adjustable fairing cone; and means for moving said diaphragmsand conoidally curved segments radially with respect to the nozzle axisin conjunction with the nozzle segments to which they are attached. 1

6. A gas turbine jet engine including, in combination, a nozzle ofvariable cross-section consisting of a succession of convergent anddivergent conical sections composed of a plurality of overlappingsegments; and means to dilate and contractsaid nozzle by moving thecomponentsegm'ents radially with respect to the nozzle'axis; a pluralityof plates within said nozzle movably attached to 'the component segments.and mechanical means to project said plates into the nozzle duct as thenozzle expands and to depress or remove them as the nozzle contracts.

References 'Cited in the file of this patent UNITED STATES PATENTS 157.526 Leggett Dec. 8, 1874 $43,182 Hunt July 23, 1895 2,390,161 MercierDec. 4, 1945 2,447,100 Stalker Aug. 17, 1948 2,462,953 Eaton et al. Mar.4, 1949 2,475,022 Gregg July 5, 1949 2,487,588 -;Price Nov. 8, 19492,509,890 Stalker May 30, 1950 2,510,506 'Lindhagen et al. June 6, 19502,569,497 Schiesel Oct. 2, 1951 2,597,253 Melchior May 20, 19522,603,062 Weiler et al. July 15, 1952 2,648,192 Lee Aug. 11, 1953FOREIGN PATENTS 617,173 Great Britain Feb. 2, 1949 922,032 France J an.20, 1947 Wed MEN

